The present invention relates generally to gas turbine engines, and, more specifically, to cooled turbine blades and stator vanes therein.
In a gas turbine engine, air is pressurized in a compressor and channeled to a combustor wherein it is mixed with fuel and ignited for generating hot combustion gases. The combustion gases flow downstream through one or more turbines which extract energy therefrom for powering the compressor and producing output power.
Turbine rotor blades and stationary nozzle vanes disposed downstream from the combustor have hollow airfoils supplied with a portion of compressed air bled from the compressor for cooling these components to effect useful lives thereof. Any air bled from the compressor necessarily is not used for producing power and correspondingly decreases the overall efficiency of the engine.
In order to increase the operating efficiency of a gas turbine engine, as represented by its thrust-to-weight ratio for example, higher turbine inlet gas temperature is required, which correspondingly requires enhanced blade and vane cooling.
Accordingly, the prior art is quite crowded with various configurations intended to maximize cooling effectiveness while minimizing the amount of cooling air bled from the compressor therefor. Typical cooling configurations include serpentine cooling passages for convection cooling the inside of blade and vane airfoils, which may be enhanced using various forms of turbulators. Internal impingement holes are also used for impingement cooling inner surfaces of the airfoils. And, film cooling holes extend through the airfoil sidewalls for providing film cooling of the external surfaces thereof.
Airfoil cooling design is rendered additionally more complex since the airfoils have a generally concave pressure side and an opposite, generally convex suction side extending axially between leading and trailing edges. The combustion gases flow over the pressure and suction sides with varying pressure and velocity distributions thereover. Accordingly, the heat load into the airfoil varies between its leading and trailing edges, and also varies from the radially inner root thereof to the radially outer tip thereof.
One consequence of the varying pressure distribution over the airfoil outer surfaces is the accommodation therefor for film cooling holes. A typical film cooling hole is inclined through the airfoil walls in the aft direction at a shallow angle to produce a thin boundary layer of cooling air downstream therefrom. The pressure of the film cooling air must necessarily be greater than the external pressure of the combustion gases to prevent backflow or ingestion of the hot combustion gases into the airfoil.
Fundamental to effective film cooling is the conventionally known blowing ratio which is the product of the density and velocity of the film cooling air relative to the product of the density and velocity of the combustion gases at the outlets of the film cooling holes. Excessive blowing ratios cause the discharged cooling air to separate or blow-off from the airfoil outer surface which degrades film cooling effectiveness. However, since various film cooling holes are fed from a common-pressure cooling air supply, providing a minimum blowing ratio for one row of commonly fed film cooling holes necessarily results in an excessive blowing ratio for the others.
Accordingly, it is desired to provide a turbine airfoil having improved film cooling notwithstanding external pressure variations therearound.